1. Field of the Invention
The present invention relates generally to rocket launch systems and, more particularly, to a low cost, expendable launch vehicle for launching consumables to low earth orbit to supply a space station or satellites.
2. Description of the Prior Art
In recent years a number of new launch vehicles (LV) have begun development with the intention of dramatically reducing launch costs. The majority of these are reusable launch vehicles (RLV) with the goal of frequent re-use in aircraft-like operations. These craft generally depend on new technology which has attendant costs and risks. The only RLV in current use, the Space Shuttle, has not lived up to promises of low operational cost and ease of use. In every instance known to the applicant, vehicles are sought with ever greater reliability and attendant cost.
It was with knowledge of the foregoing state of the technology that the present invention has been conceived and is now reduced to practice.
The present invention relates to a concept for a low-cost, moderate reliability launch vehicle to ship consumables such as water, food, oxygen, nitrogen, propellants, as well as other nonperishable items of low intrinsic value such as spare parts to low earth orbit (LEO). This concept has a goal of reducing launch costs below $1,000 per kilogram using existing technology by taking advantage of the savings inherent in a launch vehicle overall reliability of about 0.7 or less, and in the economy of scale provided by mass production of more than 1000 units annually. Infrastructure needed to support the system of the invention both at the launch site and in orbit have also been considered. Launch costs for spacecraft intended for geosynchronous orbit can be shown to be reducible by a factor of two or three depending upon the concepts and cost models used. In this regard, FIG. 1 is a graph which presents an estimate of cost of an individual launch as a function of reliability and the effective cost of lifting a payload into orbit, or payload cost. A simplistic algorithm based on empirical observations has been employed in its formulation. It is noteworthy that payload cost appears to be nearly flat in a region of reliability between 0.6 and 0.7 and remains close to a minimum value in a region of reliability between 0.5 and 0.8, indicating this to be a region of choice for reliability.
In short, the present invention flies in the face of conventional wisdom by accepting a launch vehicle of moderate reliability and accepting some amount of losses rather than one always seeking the very highest reliability humanly possible.
Preliminary studies of this new LV concept suggest that a launch success rate or overall reliability of about 0.67 is appropriate, and a launch cost of $600 K is a useful goal. For study purposes, a launch rate of 3000 annually is assumed, each carrying a payload of one metric ton. Even when the 0.67 reliability is included, payload costs in orbit would still be sufficiently below $1000/kg to permit amortization of development and other costs to establish this system while meeting the $1000/kg goal.
Uses of the payloads would include ISS (international space station) re-supply and propellant for geosynchronous orbit (GEO) and other high-value payloads launched by high-reliability LVs. Launching high-value spacecraft to LEO and then using reusable orbital maneuvering vehicles for insertion into GEO would result in substantial savings in launch costs. It is well to recognize that the environment for an orbital maneuvering vehicle is considerably less demanding than the environment for a launch vehicle. Therefore, it is appropriate to have the orbital maneuvering vehicle be reusable even though the launch vehicle is not reusable.
Costs are kept low while incurring acceptable impacts to reliability through a number of features. First, the use of a single pressure-fed main engine is assumed which provides no redundancy but avoids the thermal, contamination and other problems involved in exhaust plume interaction. This engine is ignited only once during the mission and is not re-ignited during flight minimizing ignition problems. This LV is a single-stage-to-orbit eliminating the need for staging.
A pressure-fed propellant delivery system using helium (He) gas as pressurant is assumed, eliminating the need for costly turbomachinery. This will provide an estimated regulated propellant tank pressure of 8.5 bar throughout main engine operation, to maintain engine chamber pressure of 5 bar. To aid the reader, a bar is a unit of pressure equal to one million dynes per square centimeter and thus is slightly less than one atmosphere. Gaseous nitrogen (N2) is used for a cold-gas reaction control subsystem (RCS). The RCS is used for attitude control during powered flight and He pressurant is expelled in a blow-down mode for orbital insertion and de-orbit maneuvers. A design goal is to use the RCS for all attitude control operation and avoid the need to gimbal the main engine. RCS operation is simplified by mounting, the payload aft thus center of gravity (CG) motion is favorable during powered flight. Guidance and control is simplified through the use of a GPS navigation and fiber optic gyro attitude sensing subsystem to avoid the need for costly star or earth sensors.
Use of liquid hydrogen (LH2) and liquid oxygen (LOX) for propellant is assumed. This propellant mix actually involves a lower combustion temperature than kerosene and LOX. Both of these propellants are available at relatively low cost via methanol cracking for LH2 and air liquefaction for LOX. Total propellant mass of 116.6 metric tons is estimated for a gross liftoff weight of 130 metric tons, yielding a mass when orbit is achieved of just over 13 metric tons.
The system of the invention will require supporting infrastructure both prior to launch and in orbit. Due to the relatively low reliability of this LV, many failures will occur close to liftoff, making desirable launch at sea using disposable, low-cost equipment to keep the LV afloat and erect for launch. Loading cryogenic propellant and pressurant aboard a low-reliability LV at sea necessitates procurement of sufficient quantities of remotely-operated equipment to allow for losses due to mishaps during this critical phase. In orbit, a fleet of Orbital Maneuvering Vehicles (OMV) will be used to remove payloads from the system launch vehicles and transfer the payloads to depots in higher orbits. The depots and OMVs would be launched on high-reliability LVs, expenses that will need to be amortized over several years of operations.
A primary feature, then, of the present invention is the provision of a rocket launch system utilizing an expendable launch vehicle for launching consumables to low earth orbit to supply a space station or satellites.
Another feature of the present invention is the provision of such a launch system in which a low cost launch vehicle supplies consumable items to earth orbit. In this regard, a consumable item may be defined as an item of low intrinsic value which would be expended through use in a space station or a satellite, for example, potable water in a space station or fuel in a space station or satellite. A consumable item may also be defined as an item which is inexpensively replaced in the event of a launch failure. It is the total quantity of each type of consumable item supplied to the earth orbit user over an extended span of time which is of interest, not that each individual item has the maximum or even a high probability of success in reaching the user.
Still another feature of the present invention is the provision of such a rocket launch system in which the launch vehicle need not itself rendezvous with or even approach the space station or satellite to which the consumable items are to be supplied. The launch vehicle need only be made accessible to a satellite, or orbital maneuvering vehicle, which is capable of modifying the path it follows in earth orbit, also making rendezvous and exchanging items with other orbiting craft. Such craft would include a space station or satellite to which consumable items are to be supplied and the launch vehicle. The orbital maneuvering vehicle would not be capable of entering the low altitude regions of the atmosphere which the launch vehicle traverses during launch, thus minimizing its design and operational costs. Accessibility of the launch vehicle to the orbital maneuvering vehicle means that the launch vehicle achieves orbital altitude of about 200 km (124 statute miles) or higher and does not traverse regions significantly below this altitude.
Yet another feature of the present invention is the provision of such a rocket launch system in which the launch vehicle need not maintain itself in the orbital path mentioned above. The launch vehicle and all component parts except consumable items actively removed by and under the control of the orbital maneuvering vehicle mentioned above are expected to re-enter the lower atmosphere of earth and be broken up by aerodynamic forces and heating within about one week after the 200 km (124 statute mile) altitude is achieved, thus complying with U.S. Government guidelines to avoid placing debris in orbits around earth for extended periods of time.
Still a further feature of the present invention is the provision of such a rocket launch system in which the launch vehicle can be made available for launch in sufficient quantities to assure a continuous supply of consumable items to one or more space stations and multiple satellites in earth orbit, even allowing for losses due to launch failures of the launch vehicle.
Yet a further feature of the present invention is the provision of such a rocket launch system in which the launch vehicle can be permitted to be destroyed by commands transmitted by radio by authorized persons on the ground or by commands generated by onboard computing devices evaluating input from onboard sensors in the event of a malfunction or other deviation from planned operation without delaying or affecting other launches of the launch vehicle planned for the same calendar day or any other calendar day. Onboard computing devices can be defined as a microprocessor or other computer capable of following a program written by a human operator. Sensors can be defined as suitable instruments which can determine the orientation of the launch vehicle including the direction of flight, or a pressure transducer which measures the pressure inside the fuel tank mentioned above and enables the onboard computing device to evaluate whether sufficient fuel remains in the tank to achieve the orbital conditions earlier described.
Other and further features, advantages, and benefits of the invention will become apparent in the following description taken in conjunction with the following drawings. It is to be understood that the foregoing general description and the following detailed description are exemplary and explanatory but are not to be restrictive of the invention. The accompanying drawings which are incorporated in and constitute a part of this invention, illustrate one of the embodiments of the invention, and together with the description, serve to explain the principles of the invention in general terms. Like numerals refer to like parts throughout the disclosure.